Cooling hole with shaped meter

ABSTRACT

A gas turbine engine component having a cooling passage includes a first wall defining an inlet of the cooling passage, a second wall generally opposite the first wall and defining an outlet of the cooling passage, a metering section extending downstream from the inlet, and a diffusing section extending from the metering section to the outlet. The metering section includes an upstream side and a downstream side generally opposite the upstream side. At least one of the upstream and downstream sides includes a first passage wall and a second passage wall where the first and second passage walls intersect to form a V-shape.

BACKGROUND

This invention relates generally to turbomachinery, and specifically toturbine flow path components for gas turbine engines. In particular, theinvention relates to cooling techniques for airfoils and other gasturbine engine components exposed to hot working fluid flow, including,but not limited to, rotor blades and stator vane airfoils, endwallsurfaces including platforms, blade outer air seals (shrouds) andcompressor and turbine casings, combustor liners, turbine exhaustassemblies, thrust augmentors and exhaust nozzles.

Gas turbine engines are rotary-type combustion turbine engines builtaround a power core made up of one or more compressor sections, acombustor, and one or more turbine sections arranged in flow series withan upstream inlet and downstream exhaust. The compressor section(s)compress(es) air from the inlet, which is mixed with fuel in thecombustor and ignited to generate hot combustion gas. The turbinesection(s) extract(s) energy from the expanding combustion gas, anddrive(s) the compressor section(s) via a common shaft. Expandedcombustion products are exhausted downstream, and energy is delivered inthe form of rotational energy from the shaft, reactive thrust from theexhaust, or both.

Gas turbine engines provide efficient, reliable power for a wide rangeof applications in aviation, transportation and industrial powergeneration. Small-scale gas turbine engines typically utilize aone-spool design, with co-rotating compressor and turbine sections.Larger-scale combustion turbines including jet engines and industrialgas turbines (IGTs) are generally arranged into a number of coaxiallynested spools. The spools operate at different pressures, temperaturesand spool speeds, and may rotate in different directions.

Individual compressor and turbine sections in each spool may also besubdivided into a number of stages, formed of alternating rows of rotorblade and stator vane airfoils. The airfoils are shaped to turn,accelerate and compress the working fluid flow, or to generate lift forconversion to rotational energy in the turbine.

Industrial gas turbines often utilize complex nested spoolconfigurations, and deliver power via an output shaft coupled to anelectrical generator or other load, typically using an external gearbox.In combined cycle gas turbines (CCGTs), a steam turbine or othersecondary system is used to extract additional energy from the exhaust,improving thermodynamic efficiency. Gas turbine engines are also used inmarine and land-based applications, including naval vessels, trains, andarmored vehicles, and in smaller-scale applications such as auxiliarypower units.

Aviation applications include turbojet, turbofan, turboprop andturboshaft engine designs. In turbojet engines, thrust is generatedprimarily from exhaust. Modern fixed-wing aircraft generally employturbofan and turboprop configurations, in which a low-pressure spool iscoupled to a propulsion fan or propeller. Turboshaft engines areemployed on rotary-wing aircraft, including helicopters, typically usinga reduction gearbox to control blade speed. Unducted (open rotor)turbofans and ducted propeller engines are employed in a variety ofsingle-rotor and contra-rotating designs with both forward- andaft-mounting configurations.

Aviation turbines generally utilize two- or three-spool configurations,with a corresponding number of coaxially rotating turbine and compressorsections. In two-spool designs, the high-pressure turbine drives ahigh-pressure compressor, together forming the high-pressure spool orhigh spool. The low-pressure turbine drives the low spool and fansection, or a shaft for a rotor or propeller. In three-spool engines,there is also an intermediate-pressure spool. Aviation turbines are alsoused to power auxiliary devices including electrical generators,hydraulic pumps and elements of the environmental control system, forexample using bleed air from a compressor or via an accessory gearbox.

Additional turbine engine applications and turbine engine types includeintercooled, regenerated or recuperated and variable cycle gas turbineengines, and combinations thereof. In particular, these applicationsinclude intercooled turbine engines, for example with a relativelyhigher pressure ratio, regenerated or recuperated gas turbine engines,for example with a relatively lower pressure ratio or for smaller-scaleapplications, and variable cycle gas turbine engines, for example foroperation under a range of flight conditions including subsonic,transonic and supersonic speeds. Combined intercooled andregenerated/recuperated engines with a variety of spool configurationsand traditional or variable cycle modes of operation are also used.

Turbofan engines are commonly divided into high- and low-bypassconfigurations. High-bypass turbofans generate thrust primarily from thefan, which accelerates airflow through a bypass duct oriented around theengine core. This design is common on commercial aircraft andtransports, where noise and fuel efficiency are primary concerns. Thefan rotor may also operate as a first stage compressor, or as apre-compressor stage for the low-pressure compressor or booster module.Variable-area nozzle surfaces can also be deployed to regulate thebypass pressure and improve fan performance, for example during takeoffand landing. Advanced turbofan engines may also utilize a geared fandrive mechanism to provide greater speed control, reducing noise andincreasing engine efficiency, or to increase or decrease specificthrust.

Low-bypass turbofans produce proportionally more thrust from the exhaustflow, generating greater specific thrust for use in high-performanceapplications including supersonic jet aircraft. Low-bypass turbofanengines may also include variable-area exhaust nozzles and afterburneror augmentor assemblies for flow regulation and short-term thrustenhancement. Specialized high-speed applications include continuouslyafterburning engines and hybrid turbojet/ramjet configurations.

Across these applications, turbine performance depends on the balancebetween higher pressure ratios and core gas path temperatures, whichtend to increase efficiency, and the related effects on service life andreliability due to increased stress and wear. This balance isparticularly relevant to gas turbine engine components in the hotsections of the compressor, combustor, turbine and exhaust sections,where active cooling is required to prevent damage due to high gas pathtemperatures and pressures.

SUMMARY

A gas turbine engine component with a cooling passage includes a firstwall defining an inlet of the cooling passage, a second wall generallyopposite the first wall and defining an outlet of the cooling passage, ametering section extending downstream from the inlet and a diffusingsection extending from the metering section to the outlet. The meteringsection includes an upstream side and a downstream side generallyopposite the upstream side. At least one of the upstream and downstreamsides includes a first passage wall and a second passage wall and thefirst and second passage walls intersect to form a V-shape.

A wall located in a gas turbine engine includes first and secondsurfaces and a cooling passage extending between an inlet at the firstsurface and an outlet at the second surface. The cooling passageincludes a metering section commencing at the inlet and a diffusingsection in communication with the metering section and terminating atthe outlet. The metering section includes a longitudinal first side anda longitudinal second side generally opposite the first side. The firstside includes a first passage wall and a second passage wall and thefirst and second passage walls intersect to form a vertex.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a cross-sectional view of a gas turbine engine.

FIG. 2A is a perspective view of an airfoil for the gas turbine engine,in a rotor blade configuration.

FIG. 2B is a perspective view of an airfoil for the gas turbine engine,in a stator vane configuration.

FIG. 3 is a top view of a wall having cooling passages with shapedmetering sections.

FIG. 4 is a section view of the cooling passage of FIG. 3 taken alongthe line 4-4.

FIG. 5A is a view of one embodiment of a cooling passage taken along theline 5-5 in FIG. 4.

FIG. 5B is a view of another embodiment of a cooling passage taken alongthe line 5-5 in FIG. 4.

FIG. 5C is a view of another embodiment of a cooling passage taken alongthe line 5-5 in FIG. 4.

FIG. 6 is a top view of a wall having another embodiment of a coolingpassage with a shaped metering section.

DETAILED DESCRIPTION

The present disclosure describes cooling passages with shaped meteringsections. The shaped metering sections described herein counteract thekidney vortices formed by conventional (e.g., round or oval) meteringshapes. Counteracting the formation of kidney vortices allows betterdistribution of the cooling fluid flow through the cooling passage'sdiffusing section, which is located downstream of the metering section.This results in an overall improvement of film cooling effectiveness.

FIG. 1 is a cross-sectional view of gas turbine engine 10. Gas turbineengine (or turbine engine) 10 includes a power core with compressorsection 12, combustor 14 and turbine section 16 arranged in flow seriesbetween upstream inlet 18 and downstream exhaust 20. Compressor section12 and turbine section 16 are arranged into a number of alternatingstages of rotor airfoils (or blades) 22 and stator airfoils (or vanes)24.

In the turbofan configuration of FIG. 1, propulsion fan 26 is positionedin bypass duct 28, which is coaxially oriented about the engine corealong centerline (or turbine axis) C_(L). An open-rotor propulsion stage26 can also be used, with turbine engine 10 operating as a turboprop orunducted turbofan engine. Alternatively, fan rotor 26 and bypass duct 28can be absent, with turbine engine 10 configured as a turbojet orturboshaft engine, or an industrial gas turbine.

For improved service life and reliability, components of gas turbineengine 10 are provided with an improved cooling configuration, asdescribed below. Suitable components for the cooling configurationinclude rotor airfoils 22, stator airfoils 24 and other gas turbineengine components exposed to hot gas flow including, but not limited to,platforms, shrouds, casings and other endwall surfaces in hot sectionsof compressor 12 and turbine 16, and liners, nozzles, afterburners,augmentors and other components in combustor 14 and exhaust section 20.

In the two-spool, high-bypass configuration of FIG. 1, compressorsection 12 includes low-pressure compressor (LPC) 30 and high-pressurecompressor (HPC) 32, and turbine section 16 includes high-pressureturbine (HPT) 34 and low-pressure turbine (LPT) 36. Low-pressurecompressor 30 is rotationally coupled to low-pressure turbine 36 via lowpressure (LP) shaft 38, forming the LP spool or low spool. High-pressurecompressor 32 is rotationally coupled to high-pressure turbine 34 viahigh-pressure (HP) shaft 40, forming the HP spool or high spool.

Flow F at inlet 18 divides into primary (core) flow F_(P) and secondary(bypass) flow F_(S) downstream of fan rotor 26. Fan rotor 26 acceleratessecondary flow F_(S) through bypass duct 28, with fan exit guide vanes(FEGVs) 42 to reduce swirl and improve thrust performance. In somedesigns, structural guide vanes (SGVs) 42 are used, providing combinedflow turning and load bearing capabilities.

Primary flow F_(P) is compressed in low-pressure compressor 30 andhigh-pressure compressor 32, then mixed with fuel in combustor 14 andignited to generate hot combustion gas. The combustion gas expands toprovide rotational energy in high-pressure turbine 34 and low-pressureturbine 36, which drive high-pressure compressor 32 and low-pressurecompressor 30, respectively. Expanded combustion gases exit throughexhaust section (or exhaust nozzle) 20, which can be shaped or actuatedto regulate the exhaust flow and improve thrust performance.

Low-pressure shaft 38 and high-pressure shaft 40 are mounted coaxiallyabout centerline C_(L), and rotate at different speeds. Fan rotor (orother propulsion stage) 26 is rotationally coupled to low-pressure shaft38. In advanced designs, fan drive gear system 44 is provided foradditional fan speed control, improving thrust performance andefficiency with reduced noise output.

Fan rotor 26 can also function as a first-stage compressor for gasturbine engine 10, and LPC 30 can be configured as an intermediatecompressor or booster. Alternatively, propulsion stage 26 has an openrotor design, or is absent, as described above. Gas turbine engine 10thus encompasses a wide range of different shaft, spool and turbineengine configurations, including one, two and three-spool turboprop and(high or low bypass) turbofan engines, turboshaft engines, turbojetengines, and multi-spool industrial gas turbines.

In each of these applications, turbine efficiency and performance dependon the overall pressure ratio, defined by the total pressure at inlet 18as compared to the exit pressure of compressor section 12, for exampleat the outlet of high-pressure compressor 32, entering combustor 14.Higher pressure ratios, however, also result in greater gas pathtemperatures, increasing the cooling loads on rotor airfoils 22, statorairfoils 24 and other components of gas turbine engine 10. To reduceoperating temperatures, increase service life and maintain engineefficiency, these components are provided with improved coolingconfigurations, as described below. Suitable components include, but arenot limited to, cooled gas turbine engine components in compressorsections 30 and 32, combustor 14, turbine sections 34 and 36, andexhaust section 20 of gas turbine engine 10.

FIG. 2A is a perspective view of rotor airfoil (or blade) 22 for gasturbine engine 10, as shown in FIG. 1, or for another turbomachine.Rotor airfoil 22 extends axially from leading edge 51 to trailing edge52, defining pressure surface 53 (front) and suction surface 54 (back)therebetween.

Pressure and suction surfaces 53 and 54 form the major opposing surfacesor walls of airfoil 22, extending axially between leading edge 51 andtrailing edge 52, and radially from root section 55, adjacent innerdiameter (ID) platform 56, to tip section 57, opposite ID platform 56.In some designs, tip section 57 is shrouded.

Cooling passages 60 are provided on one or more surfaces of airfoil 22,for example along leading edge 51, trailing edge 52, pressure (orgenerally concave) surface 53, or suction (or generally convex) surface54, or a combination thereof. Cooling passages 60 can also be providedon the endwall surfaces of airfoil 22, for example along ID platform 56,or on a shroud or engine casing adjacent tip section 57.

FIG. 2B is a perspective view of stator airfoil (or vane) 24 for gasturbine engine 10, as shown in FIG. 1, or for another turbomachine.Stator airfoil 24 extends axially from leading edge 61 to trailing edge62, defining pressure surface 63 (front) and suction surface 64 (back)therebetween. Pressure and suction surfaces 63 and 64 extend from inner(or root) section 65, adjacent ID platform 66, to outer (or tip) section67, adjacent outer diameter (OD) platform 68.

Cooling passages 60 are provided along one or more surfaces of airfoil24, for example leading or trailing edge 61 or 62, pressure (concave) orsuction (convex) surface 63 or 64, or a combination thereof. Coolingpassages 60 can also be provided on the endwall surfaces of airfoil 24,for example along ID platform 66 and OD platform 68.

Rotor airfoils 22 (FIG. 2A) and stator airfoils 24 (FIG. 2B) are formedof high strength, heat-resistant materials such as high-temperaturealloys and superalloys, and are provided with thermal anderosion-resistant coatings. Airfoils 22 and 24 are also provided withinternal cooling passageways and cooling passages 60 to reduce thermalfatigue and wear, and to prevent melting when exposed to hot gas flow inthe higher temperature regions of a gas turbine engine or otherturbomachine. Cooling passages 60 deliver cooling fluid (e.g., steam orair from a compressor) through the outer walls and platform structuresof airfoils 22 and 24, creating a thin layer (or film) of cooling fluidto protect the outer (gas path) surfaces from high-temperature flow.

While surface cooling extends service life and increases reliability,injecting cooling fluid into the gas path also reduces engineefficiency, and the cost in efficiency increases with increased coolingflow. Cooling passages 60 are thus provided with improved metering andinlet geometry to reduce jets and blow off, and improved diffusion andexit geometry to reduce flow separation and corner effects. Coolingpassages 60 reduce flow requirements and improve the spread of coolingfluid across the hot outer surfaces of airfoils 22 and 24, and other gasturbine engine components, so that less flow is needed for cooling andefficiency is maintained or increased.

The cooling passages described herein provide a cooling solution thatoffers improved film cooling and eliminates or reduces the flowseparation problems associated with conventional diffusion-type filmcooling passages. The shape of the cooling passage metering section ismodified to better direct cooling air to the passage's diffusing sectionand form a uniform cooling film downstream from the cooling passage. Thedescribed cooling passages provide improved film effectiveness andreduce the likelihood of film separation so that they work as intendedat high blowing ratios.

Some cooling passages include two sections: (1) a metering section at ornear the passage inlet and (2) a diffusing section at or near thepassage outlet. The metering section “meters” the flow of cooling air,regulating the velocity and quantity of air that enters through theinlet. Air flowing through the metering section enters the diffusingsection before reaching the outlet. The diffusing section causes thecooling air to expand (diffuse) so that a wider cooling film is formed.A recent trend in state of the art cooling passages has been to modifythe cooling film by changing the geometry or configuration of thediffusing section of cooling passages. While this technique has yieldedsome improvements in film cooling, it also presents additionaldifficulties. For example, some cooling passages with multi-lobeddiffusing sections can diffuse the cooling air too much at high blowingratios, spreading the cooling film too thinly so that “holes” or “gaps”in the cooling film appear. This phenomenon is called flow separation.Fluid in the hot gas path adjacent to the cooling passage can mix intothese holes or gaps in the cooling film, transferring unwanted heat tothe film cooled component and reducing cooling effectiveness.Additionally, although film cooling performance typically improves asthe blowing ratio is increased, the expansion ratio of the diffuser canbe too great, resulting in flow separation and incomplete filling of thediffuser section of the cooling passage with cooling air. In thesecircumstances, high temperature gases passing along wall surfaces canmix with the cooling air flowing within the diffuser section of thecooling passage (i.e., hot gas entrainment). The turbulent mixing thatoccurs during hot gas entrainment can adversely impact film coolingeffectiveness and performance of the diffusing section of the coolingpassage. Instead of modifying the diffusing section of the coolingpassage to reduce the incidence of flow separation and potentialentrainment of hot gas path flow (high temperature gases), whichadversely impacts overall cooling performance, geometric features areintroduced that reduce the propensity of flow separation with highlydiffused cooling passage geometries, while also mitigating the amount ofturbulent mixing that occurs between the expelled film cooling flow andthe free stream gas within the thermal boundary layer. The coolingpassages described herein contain modified metering section geometry toimprove the overall film cooling performance by improving diffusingsection fill characteristics while also reducing the amount ofdownstream film attenuation.

FIG. 3 illustrates a view of a portion of a wall of a gas turbine enginecomponent having cooling passages. Wall 100 includes inner wall surface102 and outer wall surface 104. As described in greater detail below,wall 100 is primarily metallic and outer wall surface 104 can include athermal barrier coating. Cooling passages 106 are oriented so that theirinlets are positioned on the first wall surface 102 and their outletsare positioned on outer wall surface 104. During gas turbine engineoperation, outer wall surface 104 is in proximity to high temperaturegases (e.g., combustion gases, hot air). Cooling air is delivered insidewall 100 where it exits the interior of the component through coolingpassages 106 and forms a cooling film on outer wall surface 104. Thediffusing section of cooling passage 106 can have multiple lobes to aidin the lateral diffusion of the cooling air as shown in FIG. 3. In thisembodiment, cooling passages 106 have three lobes in the diffusingsection.

As described in greater detail below, cooling air enters the meteringsection of cooling passage 106 and flows out of the diffusing section ofcooling passage 106. Cooling passages 106 can be arranged in a row onwall 100 as shown in FIG. 3 and positioned axially so that the coolingair flows in substantially the same direction longitudinally as the hightemperature gases flowing past wall 100. In this embodiment, cooling airpassing through cooling passages 106 exits cooling holes traveling insubstantially the same direction as the high temperature gases flowingalong outer wall surface 104 (represented by arrow H). Here, the linearrow of cooling passages 106 is substantially perpendicular to thedirection of flow H. In alternate embodiments, the orientation ofcooling passages 16 can be arranged on outer wall surface 104 so thatthe flow of cooling air is substantially perpendicular to the hightemperature gas flow (i.e., cooling air exits cooling passages 106radially) or at an angle between parallel and perpendicular (compoundangle). Cooling passages 106 can also be provided in a staggeredformation on wall 100. Cooling passages 106 can be located on a varietyof components that require cooling. Suitable components include, but arenot limited to, turbine vanes and blades, blade or vane platforms,shrouds, endwalls, combustors, blade outer air seals, augmentors, etc.Cooling passages 106 can be located on the pressure side or suction sideof airfoils. Cooling passages 106 can also be located on the blade tip.

FIG. 4 illustrates a sectional view of film cooling passage 106 of FIG.3 taken through the center of cooling passage 106 along the line 4-4.Cooling passage 106 includes inlet 110, metering section 112, diffusingsection 114 and outlet 116. Inlet 110 is an opening located on innerwall surface 102. Cooling air C enters cooling passage 106 through inlet110 and passes through metering section 112 and diffusing section 114before exiting cooling passage 106 at outlet 116 along outer wallsurface 104.

Metering section 112 is adjacent to and downstream from inlet 110 andcontrols (meters) the flow of cooling air through cooling passage 106.In some embodiments, metering section 112 has a substantially constantflow area from inlet 110 to diffusing section 114. Metering sections 112have a length l and hydraulic diameter d_(h). Hydraulic diameters(d_(h)) are used to describe flow in non-circular channels. In someembodiments, metering section 112 has a length l according to therelationship: d_(h)≤l≤3d_(h). That is, the length of metering section112 is between one and three times its hydraulic diameter. The length ofmetering section 112 can exceed 3d_(h), reaching upwards of 30d_(h). Insome embodiments, metering section 112 is inclined with respect to wall100 as illustrated in FIG. 4 (i.e., metering section 112 is notperpendicular to wall 100). Metering section 112 has a longitudinal axisrepresented by numeral 118. As shown in FIG. 4, metering section 112 isinclined with respect to wall 100 to include upstream side 120 anddownstream side 122. Upstream side 120 and downstream side 122 areconnected by passage side walls (shown in FIGS. 5A, 5B, and 5C) to formmetering section 112.

Diffusing section 114 is adjacent to and downstream from meteringsection 112. Cooling air C diffuses within diffusing section 114 beforeexiting cooling passage 106 along outer wall surface 104. Outer wallsurface 104 includes upstream end 124 (upstream of cooling passage 106)and downstream end 126 (downstream from cooling passage 106). Diffusingsection 114 opens along outer wall surface 104 between upstream end 124and downstream end 126. As shown in FIG. 4, cooling air C diffuses awayfrom longitudinal axis 118 in diffusing section 114 as it flows towardsoutlet 116. Diffusing section 114 can have various configurations.Diffusing section 114 can have multiple lobes as shown in FIGS. 3, 4,and 5A-5C and described in greater detail in the U.S. Pat. Nos.8,763,402; 8,683,813; and 8,584,470, each of which are incorporated byreference. For example, FIG. 3 illustrates cooling passage 106 havingdiffusing section 114 that includes three lobes while FIGS. 5A, 5B, and5C illustrate cooling passages 106 having diffusing sections 114 thateach include two lobes. In other embodiments, diffusing section 114 is amore conventional diffusing section such as those described in U.S. Pat.No. 4,197,443 or 4,684,323.

To improve the flow of cooling air C through cooling passage 106,metering section 112 does not possess the conventional circular, oblong(oval or elliptical) or racetrack (oval with two parallel sides havingstraight portions) cross-sectional geometries common in some coolingpassages. Instead, metering section 112 includes at least one sidehaving two passage walls that generally form a V-shape.

FIGS. 5A, 5B, 5C, and 6 illustrate different embodiments of coolingpassage 106 in greater detail; each embodiment has a different shapedmetering section. FIGS. 5A, 5B, and 5C show metering section 112 fromthe perspective of diffusing section 114 (i.e., the viewer is lookingstraight through metering section 112 towards inlet 110).

In the embodiment shown in FIG. 5A, upstream side 120 of meteringsection 112 contains two passage walls that form a V-shape. Upstreamside 120 includes first passage wall 128 and second passage wall 130.First passage wall 128 and second passage wall 130 intersect at vertex132 at or near the lateral (left-to-right) center of metering section112 to form a V-shape. Passage side walls 134 and 136 connect upstreamside 120 and downstream side 122. In this embodiment, first passage wall128 and second passage wall 130 are straight. Angle θ₁ represents theangle between first passage wall 128 and second passage wall 130. Insome embodiments, angle θ₁ is greater than 90 degrees. In anotherembodiment, one of the first and second passage walls is straight; theother passage wall is at least partially curved.

In the embodiment shown in FIG. 5B, downstream side 122A of meteringsection 112A contains two passage walls that form a V-shape. Downstreamside 122A includes first passage wall 128A and second passage wall 130A.First passage wall 128A and second passage wall 130A intersect at vertex132A at or near the lateral (left-to-right) center of metering section112A to form a V-shape. Passage side walls 134 and 136 connect upstreamside 120A and downstream side 122A. In this embodiment, first passagewall 128A and second passage wall 130A are straight. Angle θ₂ representsthe angle between first passage wall 128A and second passage wall 130A.In some embodiments, angle θ₂ is greater than 90 degrees. In anotherembodiment, one of the first and second passage walls is straight; theother passage wall is at least partially curved.

In the embodiment shown in FIG. 5C, both upstream side 120 anddownstream side 122A of metering section 112B each contain two passagewalls that form a V-shape. Upstream side 120 includes first passage wall128 and second passage wall 130. First passage wall 128 and secondpassage wall 130 intersect at vertex 132 at or near the lateral(left-to-right) center of metering section 112B to form a first V-shape.In this embodiment, first passage wall 128 and second passage wall 130are straight. Angle θ₁ represents the angle between first passage wall128 and second passage wall 130. In some embodiments, angle θ₁ isgreater than 90 degrees. Downstream side 122A includes third passagewall 128A and fourth passage wall 130A. Third passage wall 128A andfourth passage wall 130A intersect at vertex 132A at or near the lateral(left-to-right) center of metering section 112B to form a secondV-shape. In this embodiment, third passage wall 128A and fourth passagewall 130A are straight. Angle θ₂ represents the angle between thirdpassage wall 128A and fourth passage wall 130A. In some embodiments,angle θ₂ is greater than 90 degrees. Passage side walls 134 and 136connect upstream side 120 and downstream side 122A. As shown in FIG. 5C,metering section 112B has a cross-section with a chevron shape. In someembodiments, angle θ₂ may be equal to, greater than, or less than angleθ₁.

Metering section 112B can be configured so that first passage wall 128and third passage wall 128A are parallel and so that second passage wall130 and fourth passage wall 130A are parallel. In these cases, angles θ₁and θ₂ are equivalent. In other embodiments, such as that shown in FIG.5C, the above groups of passage walls are not parallel to one another.In these cases, angles θ₁ and θ₂ are not equivalent. In some of theembodiments where angles θ₁ and θ₂ are not equivalent, the height ofmetering section 112B is greater at its center (h₁, measured as thedistance from vertex 132 to vertex 132A) than at its ends (e.g., h₂,measured as the distance from the outer end of first passage wall 128 tothird passage wall 128A in a direction parallel to h₁). In otherembodiments h₂ is greater than h₁.

As shown in FIGS. 5A, 5B, and 5C, diffusing section 114 can containmultiple lobes. As noted above, diffusing section 114 can take a numberof different shapes, such as those described in U.S. Pat. Nos.8,763,402; 8,683,813; and 8,584,470. In some instances, diffusingsection 114 contains ridge 138 that can extend from metering section 116through diffusing section 114 to outlet 116. Ridge 138 can intersectwith vertex 132A as shown in FIGS. 5B and 5C. Thus, passage walls 128Aand 130A intersect not only with each other, but also with ridge 138.

FIG. 6 illustrates another embodiment of a cooling passage with a shapedmetering section. In this example, cooling passage 106A includes one ofthe shaped metering sections 112 described herein (e.g., meteringsection 112, 112A, 112B, etc.). The diffusing section differs from thoseof FIGS. 3, 5A, 5B, and 5C, however. Here, diffusing section 114A is ofthe type described in U.S. Pat. No. 7,328,580. As shown in FIG. 6,diffusing section 114A includes a pair of wing troughs 140 increasing inlateral width and depth along common ridge 142, which is inclined withdecreasing depth to chevron outlet 116A at outer wall surface 104.

In addition to the reduced flow separation and fill improvements notedabove, the incorporation of a shaped metering section modifies the flowstructure within the metering section of the cooling passage bygenerating counter rotating paired vortex structures having the oppositedirection of vortex structures observed in single and multi-lobediffusing sections of cooling passages with conventional cylindricalmetering shapes. The counter rotating vortices generated in shapedmetering section 112 functionally result in anti-vortices, canceling outthe vortices inherently observed in diffusing sections of coolingpassages with conventional metering shapes. The combination of the flowstructures within metering section 112 and diffusing section 114 ofcooling passage 106 results in an ejection of cooling air C thatcontains minimal or no vorticity and is laminar in nature (i.e., littledisruption). The ejection of laminar-like cooling air C reduces ormitigates the amount of turbulent mixing between the high temperaturegases flowing along wall 100 and the film coolant flow. The reducedmixing between hot and cold fluids reduces the attenuation rate of thefilm cooling boundary layer, resulting in significantly increasedadiabatic film effectiveness and film cooling performance.

Shaped metering section 112 can reduce the likelihood that cooling air Cwill diffuse in an upward (with respect to FIG. 4) direction (i.e.,forward diffusion). When forward diffusion occurs, some cooling air Cdoes not enter diffusing section 114 and jets or blows off away fromouter wall surface 104, resulting in reduced or incomplete formation ofthe film of cooling air meant to cool outer wall surface 104. Shapedmetering section 112 encourages cooling air C to diverge laterally (leftand right with respect to FIGS. 5A, 5B, and 5C) within metering section112 and once it reaches diffusing section 114. The shape of meteringsection 112 encourages cooling air C to flow into the left and rightcorners of metering section 112 and also to better attach to thesurfaces of diffusing section 114 to prevent “jet off” or “blow off” athigh blowing ratios.

The embodiments of cooling passage 106 described herein allow the use ofhigh blowing ratios of cooling air C. As the blowing ratio increases,the pressure gradient across cooling passage 106 increases. When thepressure gradient across cooling passage 106 is increased, cooling air Cis forced to fill the extremities (corners, edges, etc.) of meteringsection 112. By filling the entire metering section 112 with cooling airC, the air flow is less likely to separate one it reaches diffusingsection 114 and begins to expand. Thus, metering section 112 improvesthe filling of diffusing section 114 with cooling air C.

By encouraging lateral diffusion at high blowing ratios, diffusingsection 114 is able to provide a better film of cooling air along outerwall surface 104 and cool the gas turbine engine component. Producing abetter film of cooling air provides cooling solution flexibility. Thenumber of cooling passages 106 needed to cool the component can bereduced, the temperature of cooling air C used to cool the component canbe increased and/or the component can be exposed to higher temperatureenvironments without overheating. The cooling passages described hereinwill provide improved film cooling at any blowing ratio, but areparticularly suited for blowing ratios between about 0.5 and 10 wherethe blowing ratio (mass flux ratio) is calculated according to theequation:M=ρ _(f) V _(f) ¹/ρ_(∞) V _(∞)

Passage walls 128, 128A, 130 and 130A can also include vortex-generatingstructures such as those described in U.S. patent application Ser. No.12/157,115. Vortex-generating structures can be used to negate flowvortices that are created elsewhere in cooling hole 106 to prevent theformation of kidney vortices at outlet 116 and the unwanted entrainmentof high temperature gas into the cooling air film.

The gas turbine engine components, gas path walls and cooling passagesdescribed herein can be manufactured using one or more of a variety ofdifferent processes. These techniques can provide each cooling passagewith its own particular configuration and features, including, but notlimited to, inlet, metering, diffusion, outlet, upstream wall,downstream wall, lateral wall, longitudinal, lobe and downstream edgefeatures, as described herein. In some cases, multiple techniques can becombined to improve overall cooling performance or reproducibility, orto reduce manufacturing costs.

Suitable manufacturing techniques for forming the cooling configurationsdescribed herein include, but are not limited to, electrical dischargemachining (EDM), laser drilling, laser machining, electrical chemicalmachining (ECM), water jet machining, casting, conventional machining,additive manufacturing and combinations thereof. Electrical dischargemachining includes both machining using a shaped electrode as well asmultiple pass methods using a hollow spindle or similar electrodecomponent. Laser machining methods include, but are not limited to,material removal by ablation, trepanning and percussion laser machining.Conventional machining methods include, but are not limited to, milling,drilling and grinding. Additive manufacturing methods include, but arenot limited to, robocasting, electron-beam melting, selective lasermelting, selective laser sintering, direct metal laser sintering,directed energy deposition and electron beam freeform fabrication.

The gas flow path walls and outer surfaces of some gas turbine enginecomponents can include one or more coatings, such as bond coats, thermalbarrier coatings, abrasive coatings, abradable coatings and erosion orerosion-resistant coatings. For components having a coating, the inlet,metering portion, transition, diffusion portion and outlet coolingfeatures may be formed prior to coating application, after a firstcoating (e.g., a bond coat) is applied, or after a second or third(e.g., interlayer) coating process, or a final coating (e.g.,environmental or thermal barrier) coating process. Depending oncomponent type, cooling passage location, repair requirements and otherconsiderations, the diffusion portion and outlet features can be locatedwithin a wall or substrate, within a thermal barrier coating or othercoating layer applied to a wall or substrate, or based on combinationsthereof. The cooling geometry and other features may remain as describedabove, regardless of position relative to the wall and coating materialsor airfoil materials.

In addition, the order in which cooling features are formed and coatingsare applied may affect selection of manufacturing techniques, includingtechniques used in forming the inlet, metering portion, outlet,diffusion portion and other cooling features. For example, when athermal barrier coat or other coating is applied to the outer surface ofa gas path wall before the cooling hole or passage is produced, laserablation or laser drilling may be used. Alternatively, either laserdrilling or water jet machining may be used on a surface without athermal barrier coat. Additionally, different machining methods may bemore or less suitable for forming different features of the coolingpassage, for example, different EDM, laser machining and other machiningtechniques may be used for forming the outlet and diffusion features,and for forming the metering and inlet features.

Discussion of Possible Embodiments

The following are non-exclusive descriptions of possible embodiments ofthe present invention.

A gas turbine engine component with a cooling passage can include afirst wall defining an inlet of the cooling passage, a second wallgenerally opposite the first wall and defining an outlet of the coolingpassage, a metering section extending downstream from the inlet and adiffusing section extending from the metering section to the outlet. Themetering section can include an upstream side and a downstream sidegenerally opposite the upstream side. At least one of the upstream anddownstream sides can include a first passage wall and a second passagewall and the first and second passage walls intersect to form a V-shape.

The component of the preceding paragraph can optionally include,additionally and/or alternatively, any one or more of the followingfeatures, configurations and/or additional components.

The first passage wall and the second passage wall can be generallystraight.

The first passage wall and the second passage wall can intersect to forman angle that is greater than 90 degrees.

The first passage wall and the second passage wall can be located on theupstream side.

The first passage wall and the second passage wall can be located on thedownstream side.

The downstream side can include a third passage wall and a fourthpassage wall where the third and fourth passage walls intersect to forma second V-shape.

The first passage wall and the third passage wall can be generallyparallel and the second passage wall and the fourth passage wall can begenerally parallel.

The first V-shape can have a first angle and the second V-shape can havea second angle where the first angle and the second angle arenon-equivalent.

The first and second V-shapes can form a chevron shape having a centerand two opposite ends where the chevron shape has a greater height atits center than at either end.

The diffusing section can include multiple lobes.

The diffusing section can further include a ridge located between two ofthe multiple lobes, the first passage wall and the second passage wallcan be located on the downstream side, and the first passage wall andthe second passage wall can intersect with each other and the ridge.

The outlet at the second wall can be a chevron outlet and where thediffusing section can include a pair of wing troughs increasing inlateral width and depth along a common ridge being inclined withdecreasing depth to the chevron outlet at the second wall.

The component can be selected from the group consisting of bladeairfoils, vane airfoils, blade platforms, vane platforms, combustorliners, blade outer air seals, blade shrouds, augmentors and endwalls.

A method of forming the component can include forming the diffusingsection by electrical discharge machining first and forming the meteringsection by electrical discharge machining second.

A method of forming the component can include additively manufacturingthe component to define the cooling passage.

A wall located in a gas turbine engine can include first and secondsurfaces and a cooling passage extending between an inlet at the firstsurface and an outlet at the second surface. The cooling passage caninclude a metering section commencing at the inlet and a diffusingsection in communication with the metering section and terminating atthe outlet. The metering section can include a longitudinal first sideand a longitudinal second side generally opposite the first side. Thefirst side can include a first passage wall and a second passage walland the first and second passage walls intersect to form a vertex.

The wall of the preceding paragraph can optionally include, additionallyand/or alternatively, any one or more of the following features,configurations and/or additional components.

The first side can be the upstream side of the cooling passage.

The first side can be the downstream side of the cooling passage.

The second side can include a third passage wall and a fourth passagewall where the third and fourth passage walls can intersect to form asecond vertex.

The first vertex can form a first angle and the second vertex can form asecond angle where the first angle and the second angle arenon-equivalent.

The first and second vertices can form a chevron shape having a centerand two opposite ends where the chevron shape has a greater height atits center than at either end.

While the invention has been described with reference to an exemplaryembodiment(s), it will be understood by those skilled in the art thatvarious changes may be made and equivalents may be substituted forelements thereof without departing from the scope of the invention. Inaddition, many modifications may be made to adapt a particular situationor material to the teachings of the invention without departing from theessential scope thereof. Therefore, it is intended that the inventionnot be limited to the particular embodiment(s) disclosed, but that theinvention will include all embodiments falling within the scope of theappended claims.

The invention claimed is:
 1. A gas turbine engine component having acooling passage, the component comprising: a wall comprising: a firstwall surface defining an inlet of the cooling passage; a second wallsurface opposite the first wall surface and defining an outlet of thecooling passage; a metering section extending downstream from the inlet,the metering section comprising: an upstream side; and a downstream sideopposite the upstream side, wherein the downstream side comprises afirst passage wall and a second passage wall, and wherein the upstreamside comprises a third passage wall and a fourth passage wall, the firstpassage wall, the second passage wall, the third passage wall, and thefourth passage wall each being located in a common plane transverse to alongitudinal axis of the metering section; a diffusing section extendingfrom the metering section to the outlet, wherein the diffusing sectioncomprises multiple lobes and a ridge located between two lobes of themultiple lobes; wherein the first and second passage walls are straightwithin the common plane and extend towards the downstream side to form afirst V-shape in the common plane at the ridge; and the third and fourthpassage walls are straight within the common plane and extend towardsthe downstream side to form a second V-shape in the common plane.
 2. Thegas turbine engine component of claim 1, wherein the first passage walland the second passage wall intersect to form an angle that is greaterthan 90 degrees.
 3. The gas turbine engine component of claim 1, whereinthe first passage wall and the third passage wall are parallel, andwherein the second passage wall and the fourth passage wall areparallel.
 4. The gas turbine engine component of claim 1, wherein thefirst V-shape has a first angle and the second V-shape has a secondangle, wherein the first angle and the second angle are non-equivalent.5. The gas turbine engine component of claim 4, wherein the first andsecond V-shapes form a chevron shape having a center and two oppositeends, wherein the chevron shape has a height at the center that isgreater than a height at either of the two opposite ends.
 6. The gasturbine engine component of claim 1, wherein the first V-shape and thesecond V-shape together define a chevron exit of the metering section,and wherein the multiple lobes each increase in lateral width and depthin a direction along the ridge, the ridge being inclined relative to thesecond wall surface and having a decreasing depth to the chevron exit ofthe metering section.
 7. The gas turbine engine component of claim 1,wherein the component is selected from the group consisting of a bladeairfoil, a vane airfoil, a blade platform, a vane platform, a combustorliner, a blade outer air seal, a blade shroud, an augmentor, and anendwall.
 8. A method of forming the gas turbine engine component ofclaim 1, wherein the diffusing section is formed by electrical dischargemachining first and the metering section is formed by electricaldischarge machining second.
 9. A method of forming the gas turbineengine component of claim 1, wherein the gas turbine engine component isadditively manufactured.